EP0435770B1 - Turbomachine refroidie par air et procédé de refroidissement de cette turbomachine - Google Patents

Turbomachine refroidie par air et procédé de refroidissement de cette turbomachine Download PDF

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Publication number
EP0435770B1
EP0435770B1 EP90403776A EP90403776A EP0435770B1 EP 0435770 B1 EP0435770 B1 EP 0435770B1 EP 90403776 A EP90403776 A EP 90403776A EP 90403776 A EP90403776 A EP 90403776A EP 0435770 B1 EP0435770 B1 EP 0435770B1
Authority
EP
European Patent Office
Prior art keywords
air
turbine
compressor
high pressure
cooling
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
EP90403776A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP0435770A1 (fr
Inventor
Georges Mazeaud
François Jean-Pierre Mirville
Patrick William Roger Saillot
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Publication of EP0435770A1 publication Critical patent/EP0435770A1/fr
Application granted granted Critical
Publication of EP0435770B1 publication Critical patent/EP0435770B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F02C7/185—Cooling means for reducing the temperature of the cooling air or gas
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to gas turbines and it relates more particularly to the ventilation of high temperature turbines in aeronautical turbomachines.
  • French patent applications FR-A-2 552 164, FR-A-2 609 500 and FR-A-2 614 654 relate to cooling air accelerators which are centripetal and placed in a high compressor disc pressure and English patent application No. 2 189 845 describes a device for supplying cooling air to a turbine, device which comprises a centrifugal diffuser associated with a stage of the rotor of said turbine. See also document FR-A-1 097 374.
  • next-generation turbomachinery aerobic turbojet or ramjet or future turbojet
  • the cooling devices currently used will be clearly insufficient.
  • the high temperature of the air in the high pressure compressor of these turbomachines will not allow the direct use of this air as cooling fluid; it will be necessary to cool this air in a heat exchanger before routing it to the high pressure turbine.
  • French patent application FR-A-2 400 618 relates to a method for ventilating the high pressure turbine of a turbomachine which consists in cooling the cooling air, taken from the combustion chamber, in an air exchanger located outside the turbomachine in the annular channel where the dilution air of the fan circulates.
  • the cooled air then passes through the rear support posts of the high pressure compressor and then through a pressure reducing valve.
  • the pressure drop caused by the flow of air in the cooling circuit will not allow a sufficiently rapid circulation of the air in the blades of the high pressure turbine of the future turbomachine.
  • the object is achieved according to the invention by the fact that the method for cooling the high pressure turbine of a turbomachine comprising from upstream to downstream a high pressure compressor serving to compress air, a combustion chamber and a high pressure turbine, process according to which part of the compressed air is removed by the high pressure compressor, this air taken by cooling means is cooled and the cooled air is sent to the turbine, is characterized in that the cooled air is recompressed by the cooling means before sending it to the turbine.
  • the movable rotor forms part of the drive shaft of the high pressure compressor.
  • the mobile rotor is fixed, upstream, to the rear end of the drive shaft and, downstream, to the rotary part of said turbine.
  • the aeronautical turbomachine 1 shown in the drawing comprises from upstream to downstream: a high pressure compressor 2 of which only the top stage is shown in the drawing, a diffuser casing 3, a combustion chamber 4 and a high pressure turbine 5.
  • the high pressure compressor 2 comprises a fixed external part 6 and a mobile internal part 7 rotating around the axis 8 of the turbomachine and comprising a plurality of radial vanes 9 intended to compress the air circulating in an air stream annular 10.
  • the compressed air enters the combustion chamber 4 after being rectified by fixed vanes 11 and having passed through the diffuser casing.
  • the combustion chamber 4 has an annular shape and is delimited on the side of the axis 8 of the turbomachine by an internal wall 12 and on the external side by an external wall 13.
  • the fixed 17 and movable vanes 18 of the high pressure turbine 5 are cooled by cooling air taken out of the turbomachine 1 at the high pressure compressor 2.
  • This cooling air is directed by a first manifold 20a towards a heat exchanger 21 in which a cold fluid circulates, preferably the fuel (hydrogen or fuel).
  • a second manifold 20b After cooling in the exchanger 21, the cooled air returns to the turbomachine via a second manifold 20b.
  • the cooled air passes through at least one hollow spacer 22 which is fitted with the diffuser casing 3 and is directed for its major part P1 downstream, towards the axis 8 of the turbomachine 1, by an annular collector 23 accommodating in space annular delimited by the drive shaft 16 and the internal wall 12 of the combustion chamber 4.
  • the other part P2 of the cooled air, intended for cooling the last stage of the high pressure compressor 2 is directed towards the before the turbomachine 1 through orifices 23a provided in the front part of the annular manifold 23.
  • An axial-centrifugal compressor 24 is provided between the rear end 25 of the annular manifold 23 and the fixed 17 and movable vanes 18 of the high-pressure turbine 5.
  • This axial-centrifugal compressor 24 comprises, upstream, a plurality of stages axial 26, four in number in the drawing and, downstream, a centrifugal wheel 27. It receives the cooled air circulating in the annular collector 23, and recompresses it. At the outlet of the centrifugal wheel 27, the cooled and recompressed air supplies, in cooling air, the stationary vanes 17 through orifices 28 via a manifold 29, and the movable vanes 18 through the orifices 30.
  • the axial-centrifugal compressor 24 comprises on the one hand a fixed external casing 31 connected to the internal wall 12 of the combustion chamber 4 by means of the internal flange 32 situated at the rear of the combustion chamber 4, and integral with the internal wall 12 thereof, and, on the other hand, a movable rotor 33 connected to the drive shaft 16.
  • the external casing 31 is of the type with two half-shells to facilitate mounting of the compressor axial-centrifugal 24.
  • the movable rotor 33 is an integral part of the drive shaft 16.
  • the movable rotor 33 is fixed upstream to the rear part 34 of the drive shaft 16 and downstream to the disc. 19 on the rim of which the movable blades 18 are mounted.
  • the annular collector 23 is composed of two ferrules 35 and 36 of frustoconical shape.
  • the outer shell 35 is connected, upstream, to the internal wall 12 of the combustion chamber 4 by fixing to internal flanges 36 connecting the diffuser casing 3 and the casing of the combustion chamber 4 and, downstream, to the front part 37 of the outer casing 31 of the axial-centrifugal compressor 24.
  • the internal ferrule 36 is connected upstream, with the internal flange 38 connecting the diffuser casing 3 and the high-pressure compressor 2.
  • This internal flange 38 has the orifices 23a allowing a part of the cooling air to ventilate the blades of the high pressure compressor 2.
  • the internal ferrule 36 is connected to the external ferrule 35 in the vicinity of the axial-centrifugal compressor 24 by link arms 39 arranged opposite the blades of the first stage 26 of the axial-centrifugal compressor 24.
  • a seal 40 is provided between the internal ferrule 36 and the drive shaft 16 near the rear end 25 of the annular manifold 23.
  • the operation of the turbomachine is easily understood.
  • the hot gases circulating in the movable blades 18 of the high pressure turbine 5 drive the disc 19 in rotation, which in turn drives the drive shaft 16 of the high pressure compressor 2 and, thereby, the mobile rotor 33 of the axial-centrifugal compressor 24.
  • This method and this device make it possible to appreciably improve the circulation of the cooling air, cooled by the exchanger 21, in the vanes 17 and 18 and, thus, to decrease the temperature thereof, or allow 'Increase the temperature of the hot combustion gases and, thereby, the efficiency of the turbomachine 1, without reducing the life of the high pressure turbine.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP90403776A 1989-12-28 1990-12-27 Turbomachine refroidie par air et procédé de refroidissement de cette turbomachine Expired - Lifetime EP0435770B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8917298A FR2656657A1 (fr) 1989-12-28 1989-12-28 Turbomachine refroidie par air et procede de refroidissement de cette turbomachine.
FR8917298 1989-12-28

Publications (2)

Publication Number Publication Date
EP0435770A1 EP0435770A1 (fr) 1991-07-03
EP0435770B1 true EP0435770B1 (fr) 1993-09-15

Family

ID=9389044

Family Applications (1)

Application Number Title Priority Date Filing Date
EP90403776A Expired - Lifetime EP0435770B1 (fr) 1989-12-28 1990-12-27 Turbomachine refroidie par air et procédé de refroidissement de cette turbomachine

Country Status (5)

Country Link
US (1) US5163285A (ru)
EP (1) EP0435770B1 (ru)
JP (1) JP2559297B2 (ru)
DE (1) DE69003371T2 (ru)
FR (1) FR2656657A1 (ru)

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JP6647952B2 (ja) 2016-04-25 2020-02-14 三菱重工業株式会社 ガスタービン
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Also Published As

Publication number Publication date
FR2656657A1 (fr) 1991-07-05
JPH04209934A (ja) 1992-07-31
DE69003371D1 (de) 1993-10-21
EP0435770A1 (fr) 1991-07-03
JP2559297B2 (ja) 1996-12-04
FR2656657B1 (ru) 1994-04-22
US5163285A (en) 1992-11-17
DE69003371T2 (de) 1994-02-03

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